EP2557271B1 - Method of measuring turbine blade tip erosion - Google Patents
Method of measuring turbine blade tip erosion Download PDFInfo
- Publication number
- EP2557271B1 EP2557271B1 EP12179735.1A EP12179735A EP2557271B1 EP 2557271 B1 EP2557271 B1 EP 2557271B1 EP 12179735 A EP12179735 A EP 12179735A EP 2557271 B1 EP2557271 B1 EP 2557271B1
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- EP
- European Patent Office
- Prior art keywords
- notches
- turbine blade
- notch
- tip
- turbine
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
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- 238000000034 method Methods 0.000 title claims description 9
- 230000003628 erosive effect Effects 0.000 title description 9
- 238000012360 testing method Methods 0.000 claims description 19
- 238000004519 manufacturing process Methods 0.000 claims description 9
- 238000003754 machining Methods 0.000 claims 1
- 238000002485 combustion reaction Methods 0.000 description 5
- 238000011161 development Methods 0.000 description 5
- 239000000463 material Substances 0.000 description 3
- 238000013461 design Methods 0.000 description 2
- 239000000446 fuel Substances 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 238000001816 cooling Methods 0.000 description 1
- 230000009429 distress Effects 0.000 description 1
- 230000008030 elimination Effects 0.000 description 1
- 238000003379 elimination reaction Methods 0.000 description 1
- 238000011156 evaluation Methods 0.000 description 1
- 239000002184 metal Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000035882 stress Effects 0.000 description 1
- 230000000007 visual effect Effects 0.000 description 1
Images
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D21/00—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
- F01D21/003—Arrangements for testing or measuring
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/20—Specially-shaped blade tips to seal space between tips and stator
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/141—Shape, i.e. outer, aerodynamic form
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/50—Building or constructing in particular ways
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
- F05D2240/307—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the tip of a rotor blade
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/83—Testing, e.g. methods, components or tools therefor
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
- Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
- Y10T29/00—Metal working
- Y10T29/49—Method of mechanical manufacture
- Y10T29/49316—Impeller making
- Y10T29/49336—Blade making
Definitions
- This application relates generally to a method of measuring tip erosion of a turbine blade during development and testing of the turbine blade.
- a turbine blade can tilt or expand due to creep (because of temperature and centrifugal forces).
- creep because of temperature and centrifugal forces.
- the tip can erode over time. It is important for the turbine blade to have a proper length to reduce wear at the tip while still providing a seal between the tip and the casing.
- the gas turbine engine must be disassembled to access the hardware and the turbine blade to measure and determine any erosion, rub and tilt of the tip of the turbine blade, which is costly.
- a seal serration part at a tip of a turbine blade includes a single notch. Over time and during normal operation of the gas turbine engine, the seal serration part rubs against a case to wear the seal serration part until the notch is eventually eliminated from the tip. When it is visually determined that the notch is eliminated, this indicates that the turbine blade is approaching fracture due to creep and must be replaced.
- EP 1258598 A2 discloses a sequenced manufacturing process, in which turbine airfoils are assembled into the dovetail slots of a turbine disk. The tips of the turbine airfoils assembled in the turbine disk are then measured to determine the turbine airfoils that do not have adequate tip clearance. The tips of those turbine airfoils that do not have adequate tip clearance are then machined to obtain adequate tip clearance.
- US 7704045 discloses a prior art turbine blade with blade tip cooling notches.
- EP 2267275 discloses a prior art shroudless blade.
- the present invention provides a method of designing a turbine blade as defined in claim 1.
- the present invention further provides a turbine blade as defined in claim 6.
- the present invention further provides a gas turbine engine assembly as defined in claim 12.
- a gas turbine engine 10 such as a turbofan gas turbine engine, is circumferentially disposed about an engine centerline (or axial centerline axis 12).
- the gas turbine engine 10 includes a fan 14, a low pressure compressor 16, a high pressure compressor 18, a combustion section 20, a high pressure turbine 22 and a low pressure turbine 24.
- This application can extend to engines without a fan, and with more or fewer sections.
- Air is pulled into the gas turbine engine 10 by the fan 14 and flows through a low pressure compressor 16 and a high pressure compressor 18. Fuel is mixed with the air, and combustion occurs within the combustion section 20. Exhaust from combustion flows through a high pressure turbine 22 and a low pressure turbine 24 prior to leaving the gas turbine engine 10 through an exhaust nozzle 25.
- FIG. 2 illustrates a turbine blade 32.
- the turbine blade 32 includes a root 48 received in a rotor disk (not shown), a platform 64, an airfoil 50, and a tip 42.
- the turbine blade 32 includes a leading edge 52 and a trailing edge 54.
- the turbine blade 32 also has a pressure side 56 and a suction side 58.
- the turbine blades 32 Prior to operation of the gas turbine engine 10, there is a gap between the tip 42 of the turbine blade 32 and the casing 36.
- the turbine blades 32 expand due to heat and centrifugal forces such that the tip 42 rubs the casing 36, creating a seal.
- the tip 42 can erode and wear.
- the turbine blade 32 can also tilt, causing a different amount of erosion and wear on either the pressure side 56 or the suction side 58 of the tip 42 of the turbine blade 32.
- At least two notches 60 of known depth are formed on the tip 42 of the turbine blade 32.
- one of the at least two notches 60 is formed on the pressure side 56, and the other of the at least two notches is formed on the suction side 58 (as shown in Figure 2 ).
- the least two notches 60 are both formed on the pressure side 56 or are both formed on the suction side 58.
- a plurality of notches 60 can be formed on both the pressure side 56 and the suction side 58 (as shown in Figure 3 ).
- the at least two notches 60 function as wear indicators that indicate how much wear occurs on the tip 42 of the turbine blade 32 during testing. Based on the data obtained from the wear of the at least two notches 60, the turbine blade 32 can be designed to have a specific length based on expected expansion and wear due to creep and tilt to ensure that there is optimal contact between the turbine blade 32 and the casing 36 during operation of the gas turbine engine 10 to create a seal while reducing wear.
- the at least two notches 60 are machined. In one example, the at least two notches 60 are semi-circular in shape. The semi-circular shape minimizes stress concentration.
- notches 60 having various radii are formed on the tip 42 of the turbine blade 32.
- the notches 60 are shown for illustrative purposes only and are not shown to scale.
- a set of notches 60a and 60b is formed on the pressure side 56 and the suction side 58 of the turbine blade 32, respectively.
- Another set of notches 60c and 60d is formed closer to the trailing edge 54 on the pressure side 56 and the suction side 58 of the turbine blade 32, respectively.
- Another set of notches 60e and 60f is formed even closer to the trailing edge 54 than the set of notches 60c and 60d on the pressure side 56 and the suction side 58 of the turbine blade 32, respectively.
- the location and the radius of each of the notches 60a, 60b, 60c, 60d, 60e and 60f on the tip 42 of the turbine blade 32 are a function of design.
- the turbine blade 32 in the developmental stage has a length L that is slightly longer than that the expected length of the final design of the turbine blade 32.
- the middle notches 60c and 60d each have a radius that is equal to the amount of wear that is expected when the gas turbine engine 10 is tested. That is, once the gas turbine engine 10 is tested, it is expected that the material above the notches 60c and 60c will be rubbed away such that the bottom of the notches 60c and 60d now define the tip 42.
- the length L of the turbine blade 32 and the radius of each the notches 60c and 60d are selected such this will be the expected result. However, as explained below, this might not be the case.
- the notches 60a and 60b have a radius of 0.005 mils (0.000127 mm)
- the notches 60c and 60d have a radius of 0.010 mils (0.000254 mm)
- the notches 60e and 60f have a radius of 0.015 mils (0.000381 mm).
- the tip 42 of the turbine blade 32 can include any number of notches 60 each having any radius and the notches 60 can be placed in any location and configuration on the tip 42 of the turbine blade 32.
- the sequence and quantity of the notches 60 will be predetermined based on the needed understanding of the rub phenomenon that occurs during operating of the gas turbine engine 10 during development and testing.
- the turbine blade 32 can include a fourth set of notches 60g and 60h (shown in dashed lines in Figure 3 ) that have a radius of 0.005 mils (0.000127 mm) that is located closer to the trailing edge 54 than the notches 60e and 60f.
- the notches 60a and 60b have a radius of 0.005 mils (0.000127 mm)
- the notches 60c and 60d have a radius of 0.015 mils (0.000381 mm)
- the notches 60e and 60f have a radius of 0.010 mils (0.000254 mm)
- the notches 60g and 60h have a radius of 0.005 mils (0.000127 mm).
- the gas turbine engine 10 After the notches 60 are formed in the tip 42 of the turbine blade 32 and the gas turbine engine 10 is assembled, it is operated and tested. As the turbine blades 32 rotate and increase in temperature, they expand in length, and the tips 42 rub against the casing 36. After operation of the gas turbine engine 10 during the test ends, the turbine blades 32 cool and retract in length.
- a borescope 62 (shown schematically) is then used to view the notches 60 and determine if any of the notches 60 have be eliminated due to erosion or rub of the tip 42 against the casing 36.
- the gas turbine engine 10 includes a pre-existing hole (not shown) that is filled with a plug (not shown). The plug is removed from the pre-existing hole, and the borescope 62 is inserted into a pre-existing hole to view the tip 42 of the turbine blade 32.
- the borescope 62 is employed to view and determine how much of the tip 42 has worn away during testing of the gas turbine engine 10. As each notch 60 has a known radius, it can be determined how much of the tip 42 of the turbine blade 32 has worn away during operation by viewing the tip 42 and determining which notches 60 remain and which notches 60 have been eliminated due to wear or rub against the casing 36. From this information, the proper length of the turbine blade 32 for manufacture and actual use can be determined, and the turbine blades 32 that will be manufactured for use in actual operating gas turbine engines 10 will have this manufacturing length.
- the middle notches 60c and 60d each have a radius that is equal to the amount of wear that is expected when the gas turbine engine 10 is tested.
- the middle notches 60c and 60d have been completely eliminated during testing due to rubbing of the tip 42 with the casing 36 (which also means the notches 60a and 60b with the smaller radii have been eliminated by rubbing), but the notches 60e and 60f (which have a larger radii) remain, this indicates that 0.010 mils (0.000254 mm) of material has eroded from the airfoil 50 during the test.
- the turbine blade 32 can then be manufactured with the determined manufacturing length so that when the turbine blade 32 expands due to creep during use, the tip 42 of the turbine blade 32 contacts the casing 36 to create a proper seal while reducing wear.
- the amount of wear of the notches 60a, 60c and 60e on the pressure side 56 is compared to the amount of wear of the notches 60b, 60d and 60f on the suction side 58 of the turbine blade 32 after testing by viewing with the borescope 62. If it is viewed based on the visual appearance of the notches 60 that there is more wear on one side 56 or 58 of the turbine blade 32 than the other side 56 or 58 of the turbine blade 32 due to the elimination of more notches 60 on one side 56 or 58 of the turbine blade 32 than the other side 56 or 58 of the turbine blade, this could indicate that tilt is occurring.
- the turbine blade 32 can then be designed and manufactured to take this into account.
- the turbine blade 32 can be designed to have a length that prevents erosion and wear during actual use while still providing a seal.
- any creep and tilt can be detected and be taken into consideration when designing and determining the actual length of the turbine blades 32.
- the turbine blade 32 can be made with the proper specifications, size and length prior to manufacturing.
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Description
- This application relates generally to a method of measuring tip erosion of a turbine blade during development and testing of the turbine blade.
- During operation of a gas turbine engine, a turbine blade can tilt or expand due to creep (because of temperature and centrifugal forces). When a tip of the turbine blade rubs against a casing of the gas turbine engine, the tip can erode over time. It is important for the turbine blade to have a proper length to reduce wear at the tip while still providing a seal between the tip and the casing. During development of the gas turbine engine and the turbine blade, the gas turbine engine must be disassembled to access the hardware and the turbine blade to measure and determine any erosion, rub and tilt of the tip of the turbine blade, which is costly.
- In one prior gas turbine engine, a seal serration part at a tip of a turbine blade includes a single notch. Over time and during normal operation of the gas turbine engine, the seal serration part rubs against a case to wear the seal serration part until the notch is eventually eliminated from the tip. When it is visually determined that the notch is eliminated, this indicates that the turbine blade is approaching fracture due to creep and must be replaced.
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EP 1258598 A2 discloses a sequenced manufacturing process, in which turbine airfoils are assembled into the dovetail slots of a turbine disk. The tips of the turbine airfoils assembled in the turbine disk are then measured to determine the turbine airfoils that do not have adequate tip clearance. The tips of those turbine airfoils that do not have adequate tip clearance are then machined to obtain adequate tip clearance. -
US 7704045 discloses a prior art turbine blade with blade tip cooling notches. -
EP 2267275 discloses a prior art shroudless blade. - The present invention provides a method of designing a turbine blade as defined in claim 1.
- The present invention further provides a turbine blade as defined in claim 6.
- The present invention further provides a gas turbine engine assembly as defined in
claim 12. - These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.
-
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Figure 1 illustrates a simplified cross-sectional view of a standard gas turbine engine; -
Figure 2 illustrates a turbine blade with two notches formed on a tip; -
Figure 3 illustrates a turbine blade with multiple notches formed on the tip; and -
Figure 4 illustrates a turbine blade after operation of the gas turbine engine. - As shown in
Figure 1 , agas turbine engine 10, such as a turbofan gas turbine engine, is circumferentially disposed about an engine centerline (or axial centerline axis 12). Thegas turbine engine 10 includes afan 14, alow pressure compressor 16, ahigh pressure compressor 18, acombustion section 20, ahigh pressure turbine 22 and a low pressure turbine 24. This application can extend to engines without a fan, and with more or fewer sections. - Air is pulled into the
gas turbine engine 10 by thefan 14 and flows through alow pressure compressor 16 and ahigh pressure compressor 18. Fuel is mixed with the air, and combustion occurs within thecombustion section 20. Exhaust from combustion flows through ahigh pressure turbine 22 and a low pressure turbine 24 prior to leaving thegas turbine engine 10 through anexhaust nozzle 25. - As is known, air is compressed in the
compressors combustion section 20, and expanded in theturbines 22 and 24.Rotors 26 rotate in response to the expansion, driving thecompressors fan 14. Thecompressors compressor blades 28 and static airfoils orvanes 30. Theturbines 22 and 24 include alternating rows of metal rotating airfoils orturbine blades 32 and static airfoils orvanes 34. It should be understood that this view is included simply to provide a basic understanding of the sections in agas turbine engine 10 and not to limit the invention. This invention extends to all types of gas turbines for all types of applications, in addition to other types of turbines, such as vacuum pumps, air of gas compressors, booster pump applications, steam turbines, etc. -
Figure 2 illustrates aturbine blade 32. Theturbine blade 32 includes aroot 48 received in a rotor disk (not shown), aplatform 64, anairfoil 50, and atip 42. Theturbine blade 32 includes a leadingedge 52 and atrailing edge 54. Theturbine blade 32 also has apressure side 56 and asuction side 58. - Prior to operation of the
gas turbine engine 10, there is a gap between thetip 42 of theturbine blade 32 and thecasing 36. During operation of thegas turbine engine 10, theturbine blades 32 expand due to heat and centrifugal forces such that thetip 42 rubs thecasing 36, creating a seal. However, if theturbine blade 32 expands too much due to creep, thetip 42 can erode and wear. Theturbine blade 32 can also tilt, causing a different amount of erosion and wear on either thepressure side 56 or thesuction side 58 of thetip 42 of theturbine blade 32. - During the developmental and testing phase of the
gas turbine engine 10 and theturbine blade 32, at least twonotches 60 of known depth are formed on thetip 42 of theturbine blade 32. In one example, one of the at least twonotches 60 is formed on thepressure side 56, and the other of the at least two notches is formed on the suction side 58 (as shown inFigure 2 ). In another example outside the wording of the claims, the least twonotches 60 are both formed on thepressure side 56 or are both formed on thesuction side 58. Alternately, a plurality ofnotches 60 can be formed on both thepressure side 56 and the suction side 58 (as shown inFigure 3 ). - During development and testing of the
gas turbine engine 10, the at least twonotches 60 function as wear indicators that indicate how much wear occurs on thetip 42 of theturbine blade 32 during testing. Based on the data obtained from the wear of the at least twonotches 60, theturbine blade 32 can be designed to have a specific length based on expected expansion and wear due to creep and tilt to ensure that there is optimal contact between theturbine blade 32 and thecasing 36 during operation of thegas turbine engine 10 to create a seal while reducing wear. - In one example, the at least two
notches 60 are machined. In one example, the at least twonotches 60 are semi-circular in shape. The semi-circular shape minimizes stress concentration. - In the example shown in
Figure 3 ,notches 60 having various radii are formed on thetip 42 of theturbine blade 32. Thenotches 60 are shown for illustrative purposes only and are not shown to scale. In one example, closest to the leadingedge 52, a set ofnotches pressure side 56 and thesuction side 58 of theturbine blade 32, respectively. Another set ofnotches 60c and 60d is formed closer to thetrailing edge 54 on thepressure side 56 and thesuction side 58 of theturbine blade 32, respectively. Another set ofnotches trailing edge 54 than the set ofnotches 60c and 60d on thepressure side 56 and thesuction side 58 of theturbine blade 32, respectively. The location and the radius of each of thenotches tip 42 of theturbine blade 32 are a function of design. - The
turbine blade 32 in the developmental stage has a length L that is slightly longer than that the expected length of the final design of theturbine blade 32. In one example, themiddle notches 60c and 60d each have a radius that is equal to the amount of wear that is expected when thegas turbine engine 10 is tested. That is, once thegas turbine engine 10 is tested, it is expected that the material above the notches 60c and 60c will be rubbed away such that the bottom of thenotches 60c and 60d now define thetip 42. The length L of theturbine blade 32 and the radius of each thenotches 60c and 60d are selected such this will be the expected result. However, as explained below, this might not be the case. - In a first example, the
notches notches 60c and 60d have a radius of 0.010 mils (0.000254 mm), and thenotches tip 42 of theturbine blade 32 can include any number ofnotches 60 each having any radius and thenotches 60 can be placed in any location and configuration on thetip 42 of theturbine blade 32. The sequence and quantity of thenotches 60 will be predetermined based on the needed understanding of the rub phenomenon that occurs during operating of thegas turbine engine 10 during development and testing. - In a second example, the
turbine blade 32 can include a fourth set ofnotches Figure 3 ) that have a radius of 0.005 mils (0.000127 mm) that is located closer to the trailingedge 54 than thenotches edge 52 to the trailingedge 54, thenotches notches 60c and 60d have a radius of 0.015 mils (0.000381 mm), thenotches notches - After the
notches 60 are formed in thetip 42 of theturbine blade 32 and thegas turbine engine 10 is assembled, it is operated and tested. As theturbine blades 32 rotate and increase in temperature, they expand in length, and thetips 42 rub against thecasing 36. After operation of thegas turbine engine 10 during the test ends, theturbine blades 32 cool and retract in length. - A borescope 62 (shown schematically) is then used to view the
notches 60 and determine if any of thenotches 60 have be eliminated due to erosion or rub of thetip 42 against thecasing 36. Thegas turbine engine 10 includes a pre-existing hole (not shown) that is filled with a plug (not shown). The plug is removed from the pre-existing hole, and theborescope 62 is inserted into a pre-existing hole to view thetip 42 of theturbine blade 32. - The
borescope 62 is employed to view and determine how much of thetip 42 has worn away during testing of thegas turbine engine 10. As eachnotch 60 has a known radius, it can be determined how much of thetip 42 of theturbine blade 32 has worn away during operation by viewing thetip 42 and determining whichnotches 60 remain and whichnotches 60 have been eliminated due to wear or rub against thecasing 36. From this information, the proper length of theturbine blade 32 for manufacture and actual use can be determined, and theturbine blades 32 that will be manufactured for use in actual operatinggas turbine engines 10 will have this manufacturing length. - For example, as stated above, the
middle notches 60c and 60d each have a radius that is equal to the amount of wear that is expected when thegas turbine engine 10 is tested. Returning to the first example, as shown inFigure 4 , if themiddle notches 60c and 60d have been completely eliminated during testing due to rubbing of thetip 42 with the casing 36 (which also means thenotches notches airfoil 50 during the test. Based on this knowledge, it can be determined that theturbine blades 32 are to be manufactured with a manufacturing length that is 0.010 mils (0.000254 mm) less than the length L of theturbine blade 32 prior to the test. - In another example, if only the
notches tip 42 with thecasing 36, this indicates that 0.005 mils (0.000127 mm) of material has eroded from theairfoil 50 during the test. Based on this knowledge, it can be determined that theturbine blades 32 are to be manufactured with a manufacturing length that is 0.005 mils (0.000127 mm) less than the length L of theturbine blade 32 prior to the test. - By viewing the
notches 60 each having a known radius remaining on thetip 42 of theturbine blade 32 after the test cycle with aborescope 62, it can be determined how much of theairfoil 50 has eroded because of rub and wear with thecasing 36. Theturbine blade 32 can then be manufactured with the determined manufacturing length so that when theturbine blade 32 expands due to creep during use, thetip 42 of theturbine blade 32 contacts thecasing 36 to create a proper seal while reducing wear. - Alternately, the amount of wear of the
notches pressure side 56 is compared to the amount of wear of thenotches suction side 58 of theturbine blade 32 after testing by viewing with theborescope 62. If it is viewed based on the visual appearance of thenotches 60 that there is more wear on oneside turbine blade 32 than theother side turbine blade 32 due to the elimination ofmore notches 60 on oneside turbine blade 32 than theother side turbine blade 32 can then be designed and manufactured to take this into account. - By collecting data on erosion and wear of the
tip 42 of theturbine blade 32 during testing and determining the amount of erosion and wear to thetip 42 due to creep and/or tilt prior to manufacturing theturbine blade 32 and assembling thegas turbine engine 10 for actual use, theturbine blade 32 can be designed to have a length that prevents erosion and wear during actual use while still providing a seal. By viewing the condition and existence of thenotches 60 after testing thegas turbine engine 10 and visually evaluating their condition, presence or absence by theborescope 62 based on the known radii, any creep and tilt can be detected and be taken into consideration when designing and determining the actual length of theturbine blades 32. - By using a
borescope 62 to view the condition of thetip 42 of theturbine blade 32, it is not necessary to disassemble thegas turbine engine 10 during development and engine testing, which provides a cost saving. Evaluation and disposition of several potential distress modes (i.e., creep, erosion, and tilt) is possible without tearing down thegas turbine engine 10 and needing measuring devices. Therefore, theturbine blade 32 can be made with the proper specifications, size and length prior to manufacturing. - The foregoing description is only exemplary of the principles of the invention. Many modifications and variations are possible in light of the above teachings. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than using the example embodiments which have been specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.
Claims (12)
- A method of designing turbine blades (32) for use in operating gas turbine engines (10), the method comprising the steps of:during a developmental and testing phase of a gas turbine engine (10) and a turbine blade (32) having a pressure side (56) and a suction side (58), forming at least two notches (60) on a tip (42) of the turbine blade (32), wherein the at least two notches (60) have a different known dimension;operating the gas turbine engine including the turbine blade (32) to expand a length of the turbine blade (32) such that the tip (42) of the turbine blade (32) engages a casing (36);viewing the tip of the turbine blade (32) after the step of operating of the gas turbine engine;determining an appearance of the at least two notches (60) and the tip (42); anddetermining a proper manufacturing length of turbine blades (32) for actual use in operating gas turbine engines (10) based on the step of determining the appearance of the at least two notches (60).
- The method as recited in claim 1 wherein the at least two notches (60) have a semi-circular shape.
- The method as recited in claim 1 or 2 wherein the step of forming the at least two notches includes forming one (60a) of the at least two notches (60) on the pressure side (56) of the turbine blade (32) and forming another (60b) of the at least two notches (60) on the suction side (58) of the turbine blade (32) to determine creep.
- The method as recited in any preceding claim wherein the step of forming the at least two notches (60) includes forming at least two of the notches on the pressure side (56) of the turbine blade (32) or the suction side (58) of the turbine blade (32) to determine tilt.
- The method as recited in any preceding claim wherein the step of forming the at least two notches (60) includes machining the at least two notches (60).
- A turbine blade (32) comprising:a tip (42); andat least two notches (60) on an upper surface of the tip (42), wherein the at least two notches (60) have a different known dimension, the turbine blade (32) has a pressure side (56) and a suction side (58), the at least two notches (60) have a semi-circular shape, one (60a) of the at least two notches (60) is located on the pressure side (56) of the turbine blade (32), and another (60b) of the at least two notches is located on the suction side (58) of the turbine blade (32).
- The turbine blade as recited in claim 6 wherein at least two of the notches (60) are formed on the pressure side (56) of the turbine blade (32) or the suction side (58) of the turbine blade (32).
- The turbine blade as recited in claim 6 wherein the at least two notches comprise a first set of three notches (60a,60c,60e) located on the pressure side (56) of the turbine blade (32) and a second set of three notches (60b,60d,60f) located on the suction side (58) of the turbine blade (32), and each of the first set of three notches (60a,60c,60e) on the pressure side (56) have a different radius and each of the second set of three notches (60b,60d,60f) on the suction side (58) have a different radius.
- The turbine blade as recited in claim 8 wherein the first set of three notches (60a,60c,60e) and the second set of three notches (60b,60d,60f) each comprise a first notch (60a,60b) having a radius of 0.005 mils (0.000127 mm), a second notch (60e,60f) having a radius of 0.010 mils (0.000254 mm), and a third notch (60c,60d) having a radius of 0.015 mils (0.000381 mm), wherein the first notch (60a,60b) is located closest to a leading edge (52) of the turbine blade (32), the second notch is located between the first notch and the third notch, and the third notch is located closest to a trailing edge of the turbine blade.
- The turbine blade as recited in claim 6 wherein the at least two notches comprise a first set of four notches (60a,60c,60e,60g) located on the pressure side (56) of the turbine blade (32) and a second set of four notches (60b,60d,60f,60h) located on the suction side (58) of the turbine blade (32).
- The turbine blade as recited in claim 10 wherein the first set of four notches (60a,60c,60e,60g) and the second set of four notches (60b,60d,60f,60h) each comprise a first notch (60a,60b) having a radius of 0.005 mils (0.000127 mm), a second notch (60c,60d) having a radius of 0.015 mils (0.000381 mm), a third notch (60e,60f) having a radius of 0.010 mils (0.000254 mm), and a fourth notch (60g,60h) having a radius of 0.005 mils (0.000127 mm), wherein the first notch (60a,60b) is located closest to a leading edge (52) of the turbine blade (32), the second notch (60c,60d) is located between the first notch (60a,60b) and the third notch (60e,60f), the third notch (60e,60f) is located between the second notch (60c,60d) and the fourth notch (60g,60h), and the fourth notch (60g,60h) is located closest to a trailing edge (54) of the turbine blade (32).
- A gas turbine engine comprising:a turbine blade as recited in any of claims 6-11; anda casing (36) including a hole adapted to receive a borescope (62) to view the tip (42) of the turbine blade (32).
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US13/208,983 US9322280B2 (en) | 2011-08-12 | 2011-08-12 | Method of measuring turbine blade tip erosion |
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EP2557271A3 EP2557271A3 (en) | 2016-12-07 |
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US10526912B2 (en) | 2020-01-07 |
EP2557271A2 (en) | 2013-02-13 |
US20160230590A1 (en) | 2016-08-11 |
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US9322280B2 (en) | 2016-04-26 |
US20130039773A1 (en) | 2013-02-14 |
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